The present invention relates generally to gas turbine engines, and, more specifically, to turbine nozzles therein.
In a gas turbine engine air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases. Energy is extracted from the combustion gases in turbines, with a high pressure turbine (HPT) powering the compressor through a corresponding drive shaft therebetween, and a low pressure turbine providing output power such as powering a fan disposed upstream from the compressor in a turbofan aircraft engine application.
The typical compressor includes multiple axial stages having rotor blades decreasing in size in the downstream direction for pressurizing air in turn. The pressurized air supports combustion of the fuel injected into the combustor, and efficiency of the engine increases with the temperature of the hot combustion gases.
To withstand the hot combustion gases during operation the various combustor and turbine components subject to the heat thereof are typically made of advanced superalloy materials which maintain strength at elevated temperature and promote the durability and long life of the turbine engine. Furthermore, the various hot engine components are typically cooled using a portion of the pressurized air bled from the compressor and channeled through corresponding cooling circuits in the components.
However, any air bled from the compressor which is not utilized in the combustion process decreases the overall efficiency of the engine, and therefore must be kept to a minimum. Accordingly, durability and life of the engine must be balanced against the overall efficiency thereof.
The balance of life and efficiency typically requires various tradeoffs in the design of the various components of the engine, which components are inherently interrelated in function and performance. For example, the prior art relevant to cooling of gas turbine engine components is replete with numerous cooling circuit configurations which vary either greatly, or with minor, but significant changes.
A common example is found in the various configurations of the first stage turbine nozzle which directly receives the hottest combustion gases from the combustor. The high pressure turbine nozzle is typically formed in arcuate segments having a pair of hollow nozzle vanes fixedly joined to outer and inner bands. Pressurized cooling air is bled from the discharge end of the compressor and suitably channeled to the turbine nozzle through the outer or inner bands, or both.
Each nozzle vane has the typical airfoil configuration including a generally concave pressure side and an opposite generally convex suction side which extend in chord axially between leading and trailing edges. The profile of each vane is selected for maximizing efficiency of the nozzle in redirecting the hot combustion gases to the downstream row of HPT rotor blades which extract energy therefrom.
Each nozzle vane typically includes multiple flow passages or cavities therein through which the cooling air is channeled in various, and commonly elaborate, cooling circuits. The internal surfaces of the vanes typically include small turbulators or pins which increase the heat transfer between the internal cooling air and the hot metal sidewalls.
Perforate impingement baffles are typically used in the first stage nozzle for initially directing the cooling air in impingement against the internal surfaces of the vane prior to flow thereof laterally along the inner surfaces for discharge from various outlets formed through the vane sidewalls.
Since the vanes are directly exposed to the hottest combustion gases over their external surfaces, they include various patterns of small outlet holes therethrough which cool the sidewalls themselves, as well as providing a protective film of cooling air between the vane and the hot combustion gases. Film cooling of the external surfaces of the vane is typically provided by inclined film cooling holes extending through the pressure and suction sides of the vane for ensuring a suitable cooling air film over the external surface of the vane.
Since the leading edge of each vane is firstly exposed to the hot combustion gases which split along the opposite pressure and suction sides of the vane, the leading edge typically requires specialized cooling thereof for meeting the desired life or durability requirements of the nozzle.
The combustion gases flow differently over the concave pressure side than over the convex suction side in view of the required aerodynamic performance of those sides for proper efficiency of the turbine nozzle. Accordingly, the pressure and suction sides of each vane have different configurations of the outlet holes therein intended to correspond with the different pressure and temperature distributions of the combustion gases flowing thereover during operation.
The vane airfoil converges in the axially downstream direction to a thin trailing edge which limits the ability to introduce corresponding cooling circuits between the opposite pressure and suction sides. A row of trailing edge outlets is provided in the trailing edge where space permits for discharging some of the internal cooling air for locally cooling the trailing edge region of the vane. In turbine nozzles, the trailing edge outlets are typically located on the pressure side of the vane and terminate closely adjacent to the trailing edge.
Since gas turbine engines are designed in different configurations for different applications including military, commercial, and industrial applications for powering aircraft, ships, and electrical generators, the associated cooling configurations for the components thereof also vary significantly. In commercial aircraft engines, for example, long life or durability of the engine is desired for minimizing the periodic maintenance requirements therefor, while high efficiency is also desired for decreasing the cost of operation. Long life requires effective cooling, whereas high efficiency requires minimum bleeding of the cooling airflow.
The numerous advances in design of the modern aircraft turbofan engine results in both great efficiency and long life or durability, with actual operating experience now uncovering localized distress in hot turbine components which affects the extended life thereof. For example, the high pressure, first stage turbine nozzle that is subject to the hottest combustion gases in gas turbine engines will eventually experience oxidation and localized cracking of the vanes at the end of its life due to the repeated exposure to the hot combustion gases. The local distress regions of the nozzle vane may be at any location depending upon the specific design of the nozzle and engine, including the leading edge which first receives the hot combustion gases, or the thin trailing edge, or in between.
As indicated above, the nozzle vane cooling configurations may be specifically tailored for addressing the various cooling requirements thereof including the leading edge and trailing edge regions, but this tailoring comes with a price. A given or limited amount of cooling air is available for each nozzle vane, and that cooling air budget must be distributed over the entire vane for selectively cooling the various portions thereof. Increasing cooling air to one portion of the vane necessarily decreases cooling air to other portions for a given cooling air budget.
Furthermore, redistributing the cooling air budget in a nozzle vane correspondingly affects the overall cooling thereof, and may also affect the aerodynamic performance of the nozzle itself as the cooling air is discharged through the various outlet holes covering the vane. Yet further, the pressurized cooling air delivered to the turbine nozzle is a portion of the highest pressure compressor discharge air, which discharge air is also used for cooling the liners of the combustor itself. Another balance in the design is required for cooling the turbine nozzle as well as the combustor liners using the same source air, with corresponding limited or budgeted amounts thereof.
The great sophistication and complexity of designing modern turbofan engines is further exemplified in evaluating a pre-existing first stage HPT nozzle which has been on sale and in commercial public use for decades in the U.S. This extremely mature turbine nozzle has continually undergone small changes in the configuration thereof for further enhancing its performance and durability.
In particular, this pre-existing nozzle includes a pattern of outlet holes over both the pressure and suction sides of the nozzle vanes which use the limited budget of cooling air for effective cooling of the nozzle vanes for extended life and durability. The pattern includes rows of showerhead film cooling holes bridging the pressure and suction sides of the vane at the leading edge, and a row of trailing edge outlet slots along the pressure side. Rows of gill film cooling holes are found in the suction side downstream of the showerhead holes, and additional rows of film cooling holes are found on the pressure side downstream from the showerhead holes.
The original configuration of this pre-existing design included eight rows of showerhead holes bridging the leading edge. One row extended along the leading edge. Four rows were disposed on the pressure side aft therefrom. And, three rows were disposed on the suction side aft of the leading edge. These eight rows effectively cooled the leading edge region of the nozzle vane.
However, actual operating experience uncovered local distress or oxidation on the suction side downstream of the showerhead holes, and therefore in a modification of the original design, the aft-most end row of showerhead holes on the pressure side was moved from the pressure side to the suction side immediately downstream of the aft-most row of showerhead holes on the suction side for maintaining the original cooling air budget while addressing the local suction side distress. This modified nozzle vane has also enjoyed many years of commercial public use and success in the U.S.
However, further experience in the use of this modified nozzle design is showing local distress in the region of the pressure side leading edge where the first row of showerhead holes was removed. Furthermore, additional local distress is also being experienced on the suction side of the same vanes near the trailing edge.
This modified nozzle vane, like its parent, includes two rows of cylindrical film cooling gill holes located immediately downstream of the showerhead holes on the suction side of the vane that provide film cooling thereover to the trailing edge of the vane. Two rows of such gill holes are used for minimizing the amount of cooling air required for cooling the suction side of the vane.
In a second pre-existing first stage HPT nozzle design for a different turbofan gas turbine engine enjoying many years of successful commercial public use in the U.S., a different pattern of outlet holes is found over the pressure and suction sides of the nozzle vane, including a pair of closely spaced film cooling gill holes disposed aft of multiple rows of showerhead holes at the leading edge. This second pre-existing turbine nozzle also experienced local distress over the suction side at the trailing edge, which was addressed by spreading the aft row of gill holes downstream from the forward row of gill holes, and changing the configurations thereof from conventional cylindrical film cooling holes to conventional diffusion holes having cylindrical inlets and diverging outlets. The flow size of these gill holes remained the same for maintaining the limited cooling air budget.
There are, of course, a multitude of solutions which may be used for attempting to solve these problems of local distress at two different locations on the first pre-existing nozzle vane on the pressure side leading edge and the suction side trailing edge. However, the desire to maintain the same limited cooling air budget for turbine nozzle cooling substantially increases the difficulty of the solution.
For example, merely re-introducing the removed row of pressure side showerhead holes will correspondingly increase the cooling air requirement, which in turn can change the overall cooling performance of the nozzle vane itself, the aerodynamic performance of the nozzle, and the cooling performance of the combustion liners which also utilize compressor discharge air for cooling. Attempting to decrease the size of the showerhead holes to limit the need for additional cooling air, will correspondingly adversely affect their cooling performance at the leading edge in particular.
Furthermore, resolving the local distress on the suction side trailing edge region can also affect cooling performance of the entire nozzle, including the local distress at the pressure side leading edge.
Accordingly, it is desired to provide a turbine nozzle having an improved configuration for cooling thereof while maintaining a limited air budget therefor.